(1) Field of the Invention
The present invention lies in the narrow technical field of flight control devices for an aircraft having a rotary wing with a plurality of blades. The invention relates to a device for varying blade pitch of a lift rotor. The invention also relates to a rotary wing aircraft having such a device.
(2) Description of Related Art
A rotary wing aircraft commonly includes a main rotor for providing lift and possibly also propulsion, which rotor has a plurality of blades.
The blades of the main rotor describe a very flat cone referred to by the person skilled in the art as the “rotor cone”, with the plane of rotation of the rotor (referred to as the disk plane of the rotor and defined by the path followed by the tips of said blade) being perpendicular to the general thrust generated by the main rotor. This general thrust of the main rotor may then be resolved into a vertical lift force and a horizontal force that serves to cause the rotary wing aircraft to move in translation.
Consequently, the main rotor serves to provide the rotary wing aircraft both with lift and with propulsion.
Furthermore, by controlling the shape and the angle of inclination of the rotor cone relative to the frame of reference associated with the rotary wing aircraft, a pilot can direct said aircraft accurately.
In order to act on the rotor cone, the rotor disk is tilted relative to the drive plane of the main rotor, which drive plane is perpendicular to the mast of the main rotor (i.e. the shaft that drives the rotor in rotation). For this purpose, the angle of inclination of the rotor disk is controlled by cyclic variation of the pitch of the blade.
The rotary wing aircraft is thus provided with specific means for causing the pitch of each blade to vary, thereby varying the aerodynamic angle of attack of each blade relative to the incident air stream through which the blade passes.
In order to control the general thrust from the main rotor, both in magnitude and in direction, the pilot of the rotary wing aircraft thus generally acts on the value of the pitch angle of each blade by causing each blade to turn about its longitudinal pitch axis.
Thus, when the pilot orders a collective variation of pitch, i.e. the same variation of pitch for all of the blades, the pilot varies the magnitude of the general thrust from the main rotor so as to control the altitude and the speed of the rotary wing aircraft, with speed variation also involving a corresponding variation in drag.
In contrast, varying blade pitch collectively has no effect on the direction of that general thrust and the rotor disk remains parallel to the drive plane.
As explained above, varying the direction of the general thrust generated by the main rotor results from varying the angle of inclination (tilt) of the rotor disk, with this being done by varying the pitch of the blades not collectively, but cyclically. Under such circumstances, the pitch of a blade varies as a function of its azimuth angle, and during one complete revolution it passes between a maximum value and a minimum value, which values are obtained at opposite azimuth angles.
Cyclic variation of the pitch of the blades gives rise to cyclic variation in the thrust from the blades and thus to tilting of the rotor cone. By causing the pitch of the blades to vary cyclically, the pilot controls movement of the aircraft in translation and also controls its attitude as required for controlling it, under the effect of roll and pitching moments that are induced by the general thrust.
The device for performing these collective and cyclic variations of the pitch of the blades, and thus enabling the attitude, the lift, and the movement in translation of the aircraft to be controlled, is generally referred to by the person skilled in the art as the primary flight control or “PFC”.
Furthermore, a rotary wing aircraft in flight generates a large amount of noise and vibration, in particular as a result of rotation of the blades and the main rotor assembly. For example, while the aircraft is descending, the turbulence generated by the end of a blade lies on the path of the following blade, with which it comes into collision accompanied by a characteristic slapping sound, constituting a source of noise and vibration.
Furthermore, high levels of vibration can also be generated by the main rotor if an unbalance appears in the rotor. In general, unbalance appears on a rotating element when the center of gravity of that element is not positioned on its axis of rotation.
For an aircraft main rotor, there are various reasons why unbalance may appear, in particular a difference in the weights of the blades, a movement of a blade away from its equilibrium position, or a movement of the center of gravity of a blade.
One known solution for attenuating those noise and vibration phenomena is to increase the angle of attack of each blade periodically, e.g. in such a manner as to expel the turbulence generated by the preceding blade downwards, and also in order to compensate for the offset between the center of gravity of the rotor and its axis of rotation.
Experience shows that the frequency with which the angle of attack of the blades is varied in order to attenuate noise and vibration phenomena is a function of the number of blades of the main rotor and is proportional to the frequency of rotation of the main rotor.
Tests have also confirmed that those variations in the angle of attack of the blades need to be performed at particular frequencies that are equivalent to the product 2ω, where ω corresponds to the speed of rotation of the main rotor, and also to the products (b−1)ω, bω, and (b+1)ω, where b corresponds to the number of blades of the main rotor.
In order to attenuate noise and vibration phenomena, it is thus found that the angle of attack of the blades needs to be modified at least twice during a single rotation of the main rotor. That is why the device for producing these variations in pitch that enable noise and vibration generated by the aircraft to be attenuated is generally referred to by the person skilled in the art as “multi-cyclic control”, and is also known as “higher harmonic control” (HHC).
The prior art known to the person skilled in the art generally includes systems that separate primary flight control functions from functions of attenuating.
A swashplate device, usually referred to simply as a “swashplate”, is used for primary flight control and controls cyclic and collective pitch variation of the blades by acting on the angle of inclination of each blade. The swashplate conventionally comprises a rotary plate and a non-rotary plate. The collective and cyclic pitch controls of the pilot are connected via a linkage system to the non-rotary plate, the rotary plate being mechanically connected to each blade via a respective pitch control rod.
In order to simplify specifying components associated with the rotary plate or components associated with the non-rotary plate, two frames of reference are defined. A frame of reference referred to as the “stationary frame” is associated with the cabin of the aircraft, while a frame of reference referred to as the “rotary frame” is associated with the rotary plate and thus moves with the rotary plate relative to the stationary frame of reference.
When the pilot seeks to modify the collective pitch of the blades, the pilot acts on a control that causes the swashplate to move up or down as a whole, i.e. causes both its rotary and non-rotary plates to move together. The pitch control rods are then all moved through the same distance, which means that all of the blades change pitch by the same angle.
In contrast, when the blades are caused to vary pitch cyclically in order to cause the aircraft to go in a given direction, the swashplate is not moved vertically, but rather it is tilted relative to the mast of the main rotor. Each pitch control rod then moves in a direction and by an amount that is specific thereto, and the same goes for the pitch of the associated blade.
Although that device is effective and in widespread use, it requires considerable forces in order to maneuver the blades, particularly on heavy aircraft.
Furthermore, the presence of a swashplate penalizes the performance of the aircraft because of its large weight and dimensions that are needed as a result of the large forces to be transmitted to the blades, thereby giving rise in particular to aerodynamic disturbances that lead to an increase in drag.
It is also known that the pitch of the blades can be varied by means of flaps placed on the trailing edges of each of the blades. By varying the angle of inclination of a flap relative to the air flow, the lift generated by the blade-and-flap system is modified, and consequently, under the effect of aerodynamic forces, the pitch of each blade is modified.
Furthermore, the use of flaps makes it possible to reduce the forces that need to be deployed in order to obtain a change in the angle of attack of the blades insofar as the lift area of each flap is small compared with the lift area of the entire blade, but its effectiveness is considerable because of the speed of the air flow.
Document U.S. Pat. No. 3,095,931 describes mechanical flight controls. Flaps are situated at the trailing edge of each blade and controlled by a first linkage transmitting the movements of the swashplate to each of the flaps. The control mechanism uses a second linkage connecting together the cyclic pitch controls and the collective pitch controls at the swashplate, there being mechanical coupling between the cyclic pitch control and the collective pitch control.
Document FR 2 927 881 discloses a rotorcraft in which the rotor has as many vanes as it has blades. Each vane is associated with a single blade and is linked to the blade via a mechanical linkage. The vane, e.g. under the control of the swashplate, serves to modify the pitch angle of the associated blade. That device is particularly suitable for a two-blade rotor, with each vane then being arranged at 90° relative to the blade with which it is associated, however it can be generalized to any number of blades.
Document US 2007/0131820 seeks to eliminate the swashplate, which penalizes aircraft performance in terms of weight and drag. The flight control system described is controlled by flaps incorporated in the trailing edge of each blade, first flaps being used mainly for primary flight control and second flaps being used mainly for multi-cyclic control.
Each flap is controlled by an actuator directly incorporated in the blade. Furthermore, the second flaps, which are associated mainly with multi-cyclic control, are also capable, should the need arise, of taking over primary flight control.
The actuators incorporated in the blade may in particular be linear or rotary electric motors of the brushless direct current (BLDC) type.
Those motors and their particular application to helicopter blades are described in document WO 2008/048279.
Document U.S. Pat. No. 5,224,826 discloses a system of flaps positioned along the trailing edges of blades. Each flap is independently controlled by an electromechanical actuator, preferably of the piezoelectric type, that is incorporated in the blade. That device serves to change the pitch of the blade, and even, by virtue of the way the flaps are distributed along the blade, to modify the profile of the blade in different manners along its length.
Furthermore, document JP 2002/362496 describes a rotary wing aircraft having a mechanism for adjusting the pitch of the blades of the main rotor and a mechanism for controlling the pitch of flaps arranged on each blade. That mechanism for controlling the pitch of the flap is constituted by a first non-rotary plate fastened on a second non-rotary plate constituting the mechanism for controlling the pitch of the blades. Blade pitch control is used for maneuvering the aircraft, and flap pitch control is used for reducing the noise and vibration generated by the blades of the main rotor.
Document EP 0 936 140 discloses a system for controlling a rotor having blades with each blade having a flap positioned thereon. The pitch of the blades is controlled by means of a first rotary plate, while the pitch of the flaps is controlled by means of a second rotary plate independent of the first rotary plate, or else via actuators positioned in the rotating frame of reference.
Document US 2010/0178167 describes a method and a device for controlling the pitch of the blades of a rotor by using flaps positioned on each blade. Varying the pitch of each flap causes the pitch of the corresponding blade to be varied.
Document EP 0 729 883 discloses a multi-cyclic control system for controlling the pitch of blades of a helicopter rotor by using both cyclic actuators and multi-cyclic actuators for modifying the pitch of the blades. The cyclic actuators are located upstream from a non-rotary plate while the multi-cyclic actuators may be positioned upstream from the non-rotary plate or else they may be situated between the rotary plate and each of the blades.
The document “Reduction of helicopter blade-vortex interaction noise by active rotor control technology” published by Progress in Aerospace Sciences, Oxford, GB, Vol. 33, No. 9/10 of Sep. 1, 1997, describes in particular how to reduce the noise and vibration from the blades of a helicopter rotor by applying multi-cyclic control to the pitch of its blades. That control may be implemented in particular by using multi-cyclic actuators positioned upstream from a non-rotary plate, or indeed between a rotary plate and each of the blades. Such a system includes a control unit, a computer, and vibration sensors.
Document US 2009/0140095 discloses an electric helicopter in which the main rotor and the tail rotor are driven by respective dedicated electric motors. An electromechanical device controls pitch variation of the blades.
It can be observed that the technological background includes document U.S. Pat. No. 2,936,836, which describes an aircraft having two rotors, each having two blades. Each blade includes a flap serving to estimate and correct unbalances and differences of lift between the various blades.
Furthermore, tests have been carried out on a three-bladed main rotor concerning the configuration of three electrohydraulic actuators in the fixed frame of reference for the purpose of reducing vibration. Those actuators dedicated to multi-cyclic control act on the swashplate and are themselves controlled by means of an algorithm that analyzes vibration at certain points of the aircraft. Flight testing has demonstrated results that configuration with a three-bladed main rotor that are useful in terms of attenuating the vibration generated by the blades and the main rotor.
Finally, the ADASYS project involving Eurocopter, EADS, Daimler Chrysler Research Labs, and DLR, has shown that multi-cyclic control can be provided by a flap placed on the trailing edge of each blade of the main rotor, each flap being controlled independently by an electromechanical actuator of the piezo-ceramic type.
From the above considerations, it can be seen that using flaps on the trailing edges of blades and controlled by actuators incorporated in the blades for the purpose of primary flight control enables the performance of aircraft to be increased. Such use is accompanied by eliminating the swashplate, where its weight and its induced drag are very penalizing for the performance of such aircraft.
An object of the present invention is thus to propose an alternative solution to those designs.
It should be recalled that the blades of a rotary wing aircraft are extremely thin, i.e. the relative thickness of the aerodynamic profiles of the sections of each blade is small.
That makes it difficult to incorporate an actuator such as an electric motor or a hydraulic actuator in a blade together with its control system for modifying the angle of inclination of a flap positioned at the trailing edge of each of the blades.
It is also difficult to power such actuators, whether electrically or hydraulically, since power needs to pass from the stationary frame of reference to the rotary frame of reference, and that power must be sufficient for turning the flaps. Solutions do indeed exist for transferring such power, e.g. slip rings, however the reliability of such systems in an environment having high levels of vibration cannot always be guaranteed.
Furthermore, primary flight control must be guaranteed and safe so as to ensure the safety of the occupants of the aircraft. For that purpose, use is often made of solutions such as having multiple control circuits in order to accommodate risks of failure and in order to be able to remedy them while in flight. Consequently, when using trailing edge flaps for primary flight control, such control circuit multiplication (i.e. duplication multiplication by a number greater than two) adds additional complexity in the configuration of actuators inside the blades and in the transfer of power from the stationary frame of reference to the rotary frame of reference, as mentioned above.
In the event of a failure or a malfunction, the system as made safe by multiple control lines is capable of detecting the anomaly and of correcting it. Even if that operation takes place quickly, e.g. in less than one second, there nevertheless exists a moment during which the aircraft is not properly controlled, since the system is not operational. When using a swashplate, even if the actuators are jammed, all of the flaps and consequently all of the blades continue to operate cyclically as imposed by the swashplate. The behavior of the aircraft then remains relatively stable.
In contrast, if the swashplate is eliminated and replaced by actuators incorporated in the blades, i.e. in the rotary frame of reference, it is difficult to ensure that the flaps and consequently the blades remain synchronized over a cycle while the system is not operational. Under such circumstances, an unbalance may appear in the main rotor due to the lack of synchronization of the blades. The behavior of the aircraft can become unstable and very uncomfortable until the system has corrected the fault or the malfunction.
Consequently, in the event of a failure, the behavior of the aircraft when faced with failures is better if a swashplate is used, even if the behavior is transient and lasts only for the time needed to correct the problem.
Proper balancing of each blade is also most important for good operation of the aircraft, and the center of gravity of each blade must be in a precise position, e.g. relative to the pitch axis of the blade. Any offset in this position of the center of gravity of the blade and consequently any offset in the center of gravity of the main rotor including the blade affects both the operation of the blade and the comfort of the aircraft, considerable amounts of unbalance and vibration potentially being generated in the main rotor.
In addition, the actuator controlling the flaps, such as a piston and cylinder, has a movable portion, and as a result the center of gravity of the actuator moves a little while it is in use. Because of this, integrating actuators in blades adds elements that need to be taken into account when balancing a blade, and in particular the movement of the centers of gravity of actuators while they are in use potentially complicates such balancing, or even modifies balancing during a flight.
Furthermore, the source of power used at present for controlling the various movements of the blades is mainly of a hydraulic nature when using a swashplate, in particular for powering actuators controlling the movements of the swashplate. Although hydraulic technology has been in use for a long time and is well mastered, it nevertheless gives rise to certain constraints.
The elements needed for these functions, such as actuators and pumps are expensive and very penalizing in terms of weight. Incorporating them often requires a large amount of flexible pipework and various ducts for feeding the various elements, thus adding weight, cost, and complexity for configuration in a volume that is limited.
Furthermore, the sealing of the various hydraulic elements needs to be very thorough in an environment where there is a high level of vibration, since any reduction or loss of hydraulic pressure gives rise to a degradation or even to a loss of the flight controls of the aircraft. Hydraulic technology thus imposes constraints in terms of maintenance, and thus once more contributes to high costs in order to ensure reliability for the pumps, the actuators, and the hydraulic equipment as a whole.
Finally, protecting the environment is becoming more and more important nowadays regardless of the field of activity, whether technological or otherwise.
Hydraulic technology, in particular in terms of the recyclability of its components, and its fluids in particular, can have a negative impact on the environment.